Fuel temperature control system

Abstract

A fuel temperature control system and a method of controlling a fuel temperature for a turbine engine including supplying an aviation fuel to a fuel nozzle fluidly coupled to a combustion chamber. A fuel temperature sensor for determining at least one input parameter to define an inlet fuel temperature of the aviation fuel in the fuel nozzle. A controller for receiving the at least one input parameter and for calculating a calculated flow number in the fuel nozzle. The controller capable of comparing the calculated flow number and a reference flow number associated with a threshold during a steady state condition to determine if the aviation fuel is boiling inside the fuel nozzle.

What is claimed is: 
     
       1. A gas turbine engine for an aircraft comprising:
 a compressor section and a combustion section in serial flow arrangement along an engine centerline, the combustion section comprising: 
 a combustion chamber, 
 
       a fuel nozzle fluidly coupled to the combustion chamber for providing an aviation fuel to the combustion chamber,
 a fuel temperature control system operably coupled to the fuel nozzle, the fuel temperature control system comprising:
 a heater located upstream of the fuel nozzle, 
 a fuel temperature sensor located downstream of the heater for providing a fuel temperature output indicative of a temperature of the aviation fuel entering the fuel nozzle, 
 a first pressure sensor located downstream of the heater for providing a fuel pressure output indicative of a pressure of the aviation fuel entering the fuel nozzle, 
 a second pressure sensor located downstream of the fuel nozzle for providing a downstream pressure output indicative of a downstream pressure in the combustion chamber, and 
 a controller for receiving the fuel temperature output, fuel pressure output, and downstream pressure output as inputs to generate an in-operation fuel boiling curve, wherein the controller controls the heater according to the generated in-operation fuel boiling curve in order to maximize the temperature of the aviation fuel while minimizing fuel boiling. 
 
 
     
     
       2. The gas turbine engine of  claim 1  wherein the fuel temperature control system is integral with the fuel nozzle. 
     
     
       3. The gas turbine engine of  claim 1  wherein the fuel temperature control system is located upstream from the fuel nozzle. 
     
     
       4. A method of controlling a fuel temperature for the gas turbine engine of  claim 1 , the method comprising:
 supplying the aviation fuel to the fuel nozzle fluidly coupled to the combustion chamber; 
 determining at least one input parameter with the fuel temperature sensor to define an inlet fuel temperature of the aviation fuel in the fuel nozzle; 
 receiving at the controller the at least one input parameter; 
 calculating a calculated fuel nozzle flow number in the fuel nozzle with the controller; 
 comparing the calculated fuel nozzle flow number and a reference flow number associated with a threshold during a steady state condition to determine if the aviation fuel is boiling inside the fuel nozzle; 
 determining a maximum temperature with the calculated fuel nozzle flow number; 
 establishing an in-operation boiling curve for a remainder of a flight phase; and maintaining a temperature of the fuel through the fuel nozzle based on the in-operation boiling curve. 
 
     
     
       5. The method of  claim 4  wherein determining the at least one input parameter further comprises determining a fuel flow rate with a fuel metering unit located upstream of the fuel nozzle. 
     
     
       6. The method of  claim 5  wherein determining the at least one input parameter further comprises determining a fuel nozzle inlet pressure with a fuel pressure sensor located upstream of the fuel nozzle. 
     
     
       7. The method of  claim 6  wherein determining the at least one input parameter further comprises determining a downstream pressure with a downstream pressure sensor located proximate the combustion chamber. 
     
     
       8. The method of  claim 7  wherein determining the at least one input parameter further comprises retrieving at least one of a fuel density or a reference temperature from a look-up table. 
     
     
       9. The method of  claim 4  further comprising increasing the inlet fuel temperature to a second temperature after comparing the calculated fuel nuzzle flow number to the reference flow number. 
     
     
       10. The method of  claim 4  wherein the threshold is between 5%-20%. 
     
     
       11. The method of  claim 10  wherein the threshold is 10%. 
     
     
       12. The method of  claim 4  wherein the steady state condition includes low-altitude cruise, high-altitude cruise, or steady state climb. 
     
     
       13. A method of controlling a fuel temperature during operation of the gas turbine engine of  claim 1 , the method comprising:
 flowing a fuel through the heater to heat the fuel; 
 determining at least one input parameter of the fuel with a fuel metering unit; 
 calculating a calculated flow number associated with the fuel; 
 comparing the calculated flow number with a reference flow number associated with a threshold during a steady state condition to determine if the fuel is boiling; 
 determining a boiling temperature for the fuel; 
 generating an in-operation fuel boiling curve for the fuel based on the boiling temperature; and 
 controlling a temperature of the fuel based on the in-operation fuel boiling curve. 
 
     
     
       14. The method of  claim 13  wherein determining the at least one input parameter includes measuring a temperature of the fuel. 
     
     
       15. The method of  claim 13  wherein determining the at least one input parameter includes measuring a fuel flow rate of the fuel. 
     
     
       16. The method of  claim 13  wherein determining the at least one input parameter includes passing the fuel through an orifice sized to provide a measurable pressure drop. 
     
     
       17. The method of  claim 16  further comprising flowing the fuel downstream of the orifice through a cooler and back to the fuel tank. 
     
     
       18. The method of  claim 16  further comprising flowing the fuel downstream of the orifice through a fuel injector into a combustor of the turbine engine. 
     
     
       19. The method of  claim 13  wherein the steady state condition includes idle, takeoff, climb, or cruise. 
     
     
       20. The method of  claim 13  wherein the threshold is less than 20%.